The invention relates to a structural component and particularly a fuselage shell component for an aircraft, including a skin sheet and several stiffening profile members connected to the skin sheet at least partially by welding.
In the construction of aircraft fuselages, it has become known to connect the stiffening profile members, such as frame members and stringers, to the outer fuselage skin by means of welding, at least partially or at certain locations of the aircraft. For example, see German Patent Publication DE 196 39 667 and corresponding U.S. Pat. No. 5,841,098, or German Patent Publication DE 198 44 035. Particularly, the stringers and frame members are welded onto large format skin sheets by means of laser beam welding, so as to fabricate structural components in the form of fuselage shell components that are assembled together to form the fuselage of the aircraft.
Such fuselage shell components must have a sufficient strength and stiffness to support the ordinary operating loads applied to the aircraft fuselage, as well as extreme loads applied under unusual conditions, and a further safety margin or safety factor beyond such loads. Particularly in the future, fuselage shell components will have to satisfy a so-called xe2x80x9ctwo bay crackxe2x80x9d criteria. Namely, the fuselage shell structure will have to be able to withstand a longitudinally or circumferentially extending crack that spans or extends over two frame sections or two stringer sections (i.e. the crack extends into two bays), without resulting in a failure of the overall shell structure. In this context it is further to be assumed that the stiffening profile member at the middle of the crack is broken. Thus, the remaining structure of the fuselage shell must be able to withstand the requisite loads, without failing.
In the previously typical construction, the stiffening profile members, such as frame members and stringers, were connected to the skin sheets by riveting or adhesive bonding. Such a joining method of the stiffening profile members onto the skin sheets is disadvantageous in comparison to laser welding, because the riveting and adhesive bonding result in a greater total weight, and involve greater costs and efforts in the fabrication procedures. On the other hand, the structure resulting from such rivet connections or adhesive bonding of the stiffening profile members onto the skin sheets provides a greater residual strength and a better crack stopping characteristic (i.e. resistance to crack propagation) than a corresponding shell structure in which the stiffening profile members have been laser welded onto the skin sheets.
Particularly, with a riveted or adhesively bonded junction between the skin sheets and the stiffening profile members, a crack that initiates in the skin sheet and progresses to a location of a stiffening profile member will generally not propagate into the stiffening profile member itself, because the rivets or adhesives do not provide the necessary degree of local force coupling to transmit the crack into the stringer or frame member. Thus, while the crack in the skin sheet might propagate past the location of a stringer or frame member, it does not directly damage the associated stringer or frame member. Therefore, the respective stringer or frame member maintains its original strength and holds together the skin sheet through the rivets or adhesive on opposite sides of the crack, thereby inhibiting the propagation of the crack.
The respective stiffening profile member is able to maintain this condition for a certain number of load alternations, until the extra loading transmitted from the skin into the stiffening profile member eventually fatigues and overloads the profile member, leading to a failure of the respective stiffening profile member. At that point, the fuselage skin and the affected stiffening profile member will fail, typically in a sudden rupturing manner, which leads to a failure of the fuselage shell structure. However, the fact that the stiffening profile member maintains its integrity and load-carrying ability even after a crack has formed in the adjoining skin sheet, generally allows the aircraft to fly safely to a landing, whereupon the crack defect in the skin sheet can be detected and repaired.
The above described advantageous property of crack propagation resistance or inhibition is not generally achieved by fuselage shell structures in which the stiffening profile members are welded onto the skin sheets. Namely, with such a welded junction, any crack that forms in a skin sheet and propagates to the junction of a stiffening profile member will directly propagate through the welded joint into the stiffening profile member, where the crack will then propagate further into or even entirely through the stiffening profile member. Since there is no effective interruption between the skin sheet and the stiffening profile member, there is no xe2x80x9ccrack stoppingxe2x80x9d effect which would prevent the crack from propagating into the respective stringer or frame member. As a result, any crack in the skin sheet will readily propagate through the stringers and frame members as well, which leads to a significantly lower residual or remaining strength of the overall fuselage shell structure upon the occurrence of such a crack. Namely, once such crack forms, it will readily propagate through both the skin and the stiffening profile members, and there is no structural component remaining to hold together the fuselage shell at the location of the crack, thus leading to a failure of the overall shell structure.
In view of the above, the shell structure components would have to be thickened and thereby strengthened in areas of the aircraft fuselage in which the post-crack residual strength is the predominant design criterium, in order to achieve an adequate residual strength in such areas. These areas especially include the sides and the upper or top portion of the fuselage, since these areas are especially subjected to tension loads during operation, with a consequent tendency toward crack opening and propagation. Such thickening of the fuselage shell in these areas would, however, lead to an unacceptable increase in the overall weight of the fuselage. For these reasons, prior aircraft fuselages have not used welded stringers in these areas at the sides and top of the fuselage, but instead have only used welded stringers, for example, in the bottom or belly of the fuselage, while using riveted or adhesively bonded stringers on the sides and top of the fuselage.
German Patent DE 199 24 909 has further disclosed a fuselage shell component in which each stiffening profile member includes an integral thickening at a location adjacent to the base or root of the profile member at which the profile member is welded onto a skin sheet. The ratio of the thickness of this thickening or protruding portion of the profile member relative to the thickness of the root of the profile member that is welded onto the skin sheet is at least two to one. The protruding portion or thickening is an integrally formed portion of the same material as the rest of the profile member. The object of this thickened portion or protrusion is to stop the propagation of any crack that might progress from the skin sheet through the welded junction into the base or root of the profile member. Thus, even if the crack propagates into the base or root of the profile member, it shall not propagate further beyond the thickened protrusion into the rest of the profile member. This provides a crack propagation stopping characteristic as well as an improved residual strength of the fuselage shell structure after a crack has formed in the skin. While such an integral protrusion or thickening of the stiffening profile member aims to provide a certain degree of crack stopping performance, it has been found that further improvements are possible.
In view of the above, it is an object of the invention to provide a shell structural component and particularly an aircraft fuselage shell component that has an increased residual strength after a crack has formed in the skin thereof, under consideration of a minimum structural weight of the shell component. It is a further object of the invention to provide a shell component with stiffening profile members welded onto the skin thereof that is suitable for use in all locations of the aircraft fuselage, including the sides and the top of the fuselage shell. The invention further aims to avoid or overcome the disadvantages of the prior art, and to achieve additional advantages, as apparent from the present specification.
The above objects have been achieved according to the invention in a structural shell component for an aircraft fuselage, comprising a skin sheet as well as plural stiffening profile members such as stringers or frame members, whereby the stiffening profile members are at least partially joined to the skin sheet by means of welding. Particularly according to the invention, non-unitary or non-integral strengthening elements are arranged on and secured to the stiffening profile members, before the stiffening profile members are welded onto the skin sheet. Each of these strengthening elements is a separate, non-integral component relative to the stiffening profile member onto which it is secured, and preferably consists of a different material than the stiffening profile member. In this manner, the strengthening and crack stopping effect of the strengthening elements can be optimized or maximized, without unacceptably increasing the weight or the costs of the finished structure. In other words, the strengthening members can consist of material that is stronger and lighter, but more costly, than that of the stiffening profile members, for example.
Further, preferably, the strengthening elements are secured to the stiffening profile members by a non-integral connection method, or joining method, such as riveting or adhesive bonding. Such a non-integral connection provides the crack stopping interruption that is necessary for preventing a crack from propagating into the strengthening elements or thereby also further into the stiffening profile members. Namely, if a crack propagates through the welded joint from the skin sheet into the stiffening profile member, it will not further propagate through the riveted or adhesive joint into the strengthening element or elements. Thereby the strengthening element or elements will maintain its strength intact and hold together the stiffening profile member at the location of the crack, which will inhibit the further propagation of the crack in the stiffening profile member.
According to preferred embodiments of the invention, the strengthening elements may be in the form of doubling or reinforcing members that are secured to the webs of the respective stiffening profile members, or alternatively the strengthening elements may comprise tension bands or cables that are secured to the stiffening profile members so as to extend along the respective length thereof.
The inventive shell structure achieves the advantage that the residual or remaining strength of the shell structure after a crack has formed in the skin thereof, is sufficient so that the welded shell component can also be used in the side and top areas of an aircraft fuselage. Thus, it becomes possible to use such welded fuselage shell components for the entire fuselage of an aircraft, so that the use of riveted and adhesively bonded joints between the stringers or frame members and the fuselage skin can be avoided, in all areas of the aircraft rather than only the lower belly of the fuselage. In this manner, the overall production effort, costs and structural weight can be significantly reduced in comparison to the use of conventional adhesively bonded or riveted shell components.
Moreover, the inventive provision of a two-part, non-integral structure of separate strengthening elements secured onto the stiffening profile members achieves additional advantages over the prior art provision of an integral thickening or protrusion adjacent to the base of each stiffening profile member. Namely, the use of separate or discrete strengthening elements allows the strengthening elements to be made of a different material than that of the stiffening profile members, which allows a greater strength and a greater strength-to-weight ratio to be achieved, without excessively increasing the costs. Also, the strengthening elements may be arranged with an oriented characteristic, for example, in the manner of a tension band or cable that extends along the length of the respective stiffening profile member, so as to exert its strongest retaining forces in a direction that is most effective for holding together the respective stiffening profile member across a crack, in the event a crack should propagate into the stiffening profile member. The non-integral joint between the respective strengthening element and the stiffening profile member provides better crack stopping isolation to prevent the further propagation of a crack through or beyond such a joint.
The inventive structure has thus solved or overcome all of the prior art disadvantages of welded shell components, including those that arise when using stiffening profile members having thickened portions along the roots or bases thereof. If a primary crack develops in the fuselage skin, this crack might propagate through the welded joint into the stiffening profile members, but there the crack propagation will be delayed or entirely stopped by the strengthening elements arranged according to the invention on the stiffening profile members. This in turn has the effect of stopping or hindering the propagation of the crack further in the fuselage skin. The structure of interconnected frame members and stringers remains substantially intact and maintains its strength, so that the residual or remaining strength after the initiation of a crack in the fuselage shell is increased.